Multiple piece engine component

ABSTRACT

One exemplary embodiment of this disclosure relates to a gas turbine engine, including a component having a first portion formed using one of a casting and a forging process, and a second portion formed using an additive manufacturing process.

BACKGROUND

Gas turbine engines typically include a compressor section, a combustorsection and a turbine section. During operation, air is pressurized inthe compressor section and is mixed with fuel and burned in thecombustor section to generate hot combustion gases. The hot combustiongases are communicated through the turbine section, which extractsenergy from the hot combustion gases to power the compressor section andother gas turbine engine loads.

Both the compressor and turbine sections may include alternating arraysof rotating blades and stationary vanes that extend into the core flowpath of the gas turbine engine. For example, in the turbine section,turbine blades rotate and extract energy from the hot combustion gasesthat are communicated along the core flow path of the gas turbineengine. Turbine blades are known to include an airfoil section, overwhich the hot combustion gases flow, and a root attached to a rotatabledisc. Turbine blades are typically cast such that the airfoil sectionand the root are integrally formed as a single-piece structure.

SUMMARY

One exemplary embodiment of this disclosure relates to a gas turbineengine, including a component having a first portion formed using one ofa casting and a forging process, and a second portion formed using anadditive manufacturing process.

In a further embodiment of any of the above, the component includes athird portion formed using an additive manufacturing process.

In a further embodiment of any of the above, the second portion and thethird portion provide pressure and suction side walls of the component.

In a further embodiment of any of the above, the first portion includesa root, platform, and at least one radial support projecting from theplatform.

In a further embodiment of any of the above, the root, platform, and atleast one radial support are integrally formed of one of a singlecrystal, directionally solidified, and an equiax alloy.

In a further embodiment of any of the above, the at least one radialsupport includes at least one rib projecting into a corresponding slotformed in one of the second portion and the third portion.

In a further embodiment of any of the above, the at least one radialsupport includes a plurality of radial supports, each of the radialsupports including a first rib and a second rib projecting into slotsformed in the second portion and the third portion.

In a further embodiment of any of the above, the at least one radialsupport provides a mate face corresponding to a mate face of one of thesecond portion and the third portion.

In a further embodiment of any of the above, the second portion and thethird portion are joined to the at least one radial support by one ofwelding, brazing, diffusion bonding, and gluing.

In a further embodiment of any of the above, the second portion and thethird portion include microchannels formed therein.

In a further embodiment of any of the above, the component is one of arotor blade and a stator vane.

Another exemplary embodiment of this disclosure relates to a componentfor a gas turbine engine. The component includes a platform, and anairfoil section including a pressure side wall and a suction side wall.The platform is formed using one of a casting and a forging process, andthe pressure and suction side walls are formed using an additivemanufacturing process.

In a further embodiment of any of the above, the component includes atleast one radial support projecting from the platform, the at least oneradial support formed integrally with the platform during the castingprocess.

In a further embodiment of any of the above, the at least one radialsupport provides a joining interface with one of the pressure side walland the suction side wall.

In a further embodiment of any of the above, the at least one radialsupport includes a rib projecting into a corresponding slot in one ofthe pressure side wall and the suction side wall.

Another exemplary embodiment of this disclosure relates to a method offorming a component. The method includes forming a first portion of thecomponent using one of a casting and a forging process, additivelymanufacturing a second portion of the component, and joining the secondportion to the first portion.

In a further embodiment of any of the above, the method includesadditively manufacturing a third portion of the component, and joiningthe third portion to the first portion.

In a further embodiment of any of the above, the joining step includesone of welding, brazing, and gluing.

In a further embodiment of any of the above, the additive manufacturingstep includes one of a direct metal laser sintering (DMLS) process, anelectron beam melting (EBM) process, electron beam wire deposition(EBWD) process, a laser powder deposition (LPD) process, and a laserpowder plasma spray (LPPS) process.

In a further embodiment of any of the above, the additive manufacturingstep includes selectively melting a powdered metal, and the powderedmetal is one of (1) a titanium alloy, (2) tungsten alloy, (3) nickelalloy, (4) cobalt alloy, (5) steel alloy, and (6) a molybdenum alloy.

The embodiments, examples and alternatives of the preceding paragraphs,the claims, or the following description and drawings, including any oftheir various aspects or respective individual features, may be takenindependently or in any combination. Features described in connectionwith one embodiment are applicable to all embodiments, unless suchfeatures are incompatible.

BRIEF DESCRIPTION OF THE DRAWINGS

The drawings can be briefly described as follows:

FIG. 1 schematically illustrates a gas turbine engine.

FIG. 2 is a side view of an engine component according to thisdisclosure.

FIG. 3 is a perspective, exploded view of the engine component of FIG.2.

FIG. 4 is a sectional view of the engine component of FIG. 2, takenalong line 4-4.

FIG. 5 is a sectional view of the engine component of FIG. 2, takenalong line 5-5.

FIG. 6 schematically illustrates an example method according to thisdisclosure.

FIG. 7 schematically illustrates an example laser powder plasma spray(LPPS) process.

FIG. 8A illustrates another example engine component according to thisdisclosure.

FIG. 8B is an exploded view of the engine component of FIG. 8A.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates an example gas turbine engine 20 thatincludes a fan section 22, a compressor section 24, a combustor section26 and a turbine section 28. Alternative engines might include anaugmenter section (not shown) among other systems or features. The fansection 22 drives air along a bypass flow path B while the compressorsection 24 draws air in along a core airflow C along a core flow pathwhere air is compressed and communicated to a combustor section 26. Inthe combustor section 26, air is mixed with fuel and ignited to generatea high pressure exhaust gas stream that expands through the turbinesection 28 where energy is extracted and utilized to drive the fansection 22 and the compressor section 24.

Although the disclosed non-limiting embodiment depicts a turbofan gasturbine engine, it should be understood that the concepts describedherein are not limited to use with turbofans as the teachings may beapplied to other types of turbine engines; for example a turbine engineincluding a three-spool architecture in which three spoolsconcentrically rotate about a common axis and where a low spool enablesa low pressure turbine to drive a fan via a gearbox, an intermediatespool that enables an intermediate pressure turbine to drive a firstcompressor of the compressor section, and a high spool that enables ahigh pressure turbine to drive a high pressure compressor of thecompressor section.

The example engine 20 generally includes a low speed spool 30 and a highspeed spool 32 mounted for rotation about an engine central longitudinalaxis A relative to an engine static structure 36 via several bearingsystems 38. It should be understood that various bearing systems 38 atvarious locations may alternatively or additionally be provided.

The low speed spool 30 generally includes an inner shaft 40 thatconnects a fan 42 and a low pressure (or first) compressor section 44 toa low pressure (or first) turbine section 46. The inner shaft 40 drivesthe fan 42 through a speed change device, such as a geared architecture48, to drive the fan 42 at a lower speed than the low speed spool 30.The high-speed spool 32 includes an outer shaft 50 that interconnects ahigh pressure (or second) compressor section 52 and a high pressure (orsecond) turbine section 54. The inner shaft 40 and the outer shaft 50are concentric and rotate via the bearing systems 38 about the enginecentral longitudinal axis A.

A combustor 56 is arranged between the high pressure compressor 52 andthe high pressure turbine 54. In one example, the high pressure turbine54 includes at least two stages to provide a double stage high pressureturbine 54. In another example, the high pressure turbine 54 includesonly a single stage. As used herein, a “high pressure” compressor orturbine experiences a higher pressure than a corresponding “lowpressure” compressor or turbine.

The example low pressure turbine 46 has a pressure ratio that is greaterthan about five (5). The pressure ratio of the example low pressureturbine 46 is measured prior to an inlet of the low pressure turbine 46as related to the pressure measured at the outlet of the low pressureturbine 46 prior to an exhaust nozzle.

A mid-turbine frame 57 of the engine static structure 36 is arrangedgenerally between the high pressure turbine 54 and the low pressureturbine 46. The mid-turbine frame 57 further supports bearing systems 38in the turbine section 28 as well as setting airflow entering the lowpressure turbine 46.

The core airflow C is compressed by the low pressure compressor 44 thenby the high pressure compressor 52 mixed with fuel and ignited in thecombustor 56 to produce high speed exhaust gases that are then expandedthrough the high pressure turbine 54 and low pressure turbine 46. Themid-turbine frame 57 includes vanes 59, which are in the core airflowpath and function as an inlet guide vane for the low pressure turbine46. Utilizing the vane 59 of the mid-turbine frame 57 as the inlet guidevane for low pressure turbine 46 decreases the length of the lowpressure turbine 46 without increasing the axial length of themid-turbine frame 57. Reducing or eliminating the number of vanes in thelow pressure turbine 46 shortens the axial length of the turbine section28. Thus, the compactness of the gas turbine engine 20 is increased anda higher power density may be achieved.

FIG. 2 is a side view of an engine component 62 according to thisdisclosure. For exemplary purposes, the illustrated engine component 62is a turbine blade. It should be understood that this disclosure extendsto other engine components, such as compressor blades, stator vanes(e.g., FIGS. 8A-8B), and fan blades, as non-limiting examples.

The example engine component 62 includes an airfoil section 64, a root66, and a platform 68. In this example, the root 66 includes a fir treeconfiguration. Other roots, such as dovetail roots, come within thescope of this disclosure, however.

The airfoil section 64 includes a pressure side wall 70 and a suctionside wall 72, each of which extend between a leading edge 74 and atrailing edge 76. The pressure side wall 70 and the suction side wall 72extend radially from the platform 68 to a radially outer blade tip 78.The term “radially,” as used herein refers to the radial direction Z,which is normal to the engine central longitudinal axis A, and is usedfor purposes of explaining the relative location of the illustratedcomponents without being otherwise limiting.

In one example of this disclosure, the engine component 62 is a multiplepiece engine component. For example, as illustrated in FIG. 3, thepressure side wall 70 and the suction side wall 72 can be formedseparately from the root 66 and the platform 68 of the engine component62. In other words, the exposed airfoil flowpath section 64 is aseparate structure from the root 66 and platform 68.

In one example, the root 66, the platform 68, and a plurality of radialsupports 80A-80D are integrally formed using a casting process, asexplained in detail below relative to FIG. 6. In another example, aforging process is used. While in this example four radial supports80A-80D are illustrated, it should be understood that there could be anynumber of radial supports.

The pressure side wall 70 and the suction side wall 72 are separatelyformed using an additive manufacturing process, again, as will beexplained below relative to FIG. 6. After forming, the pressure sidewall 70 and the suction side wall 72 are then joined to the radialsupports 80A-80D to provide the airfoil section 64, as will be describedin detail below.

Referring to FIG. 3, the pressure side wall 70 includes an outer surface82 providing a portion of the outer contour of the airfoil section 64,and an inner surface 84 facing the radial supports 80A-80D. In oneexample, the pressure side wall 70 includes a plurality of radiallyextending slots 86A-86D.

In this example, the slots 86A-86D extend generally parallel to oneanother, and have a longitudinal dimension extending generally in theradial direction Z. The slots 86A-86D correspond to a plurality ofpressure side ribs 88A-88D extending from the radial supports 80A-80D ina direction perpendicular to the radial direction Z.

The suction side wall 72 likewise includes an outer surface 90, an innersurface 92, and a plurality of slots 94A-94D which correspond to aplurality of suction side ribs 96A-96D extending from a suction side ofthe radial supports 80A-80D.

The slots 86A-86D, 94A-94D and ribs 88A-88D, 96A-96D facilitatealignment of the pressure side wall 70 and the suction side wall 72relative to the radial supports 80A-80D, which may increase the ease ofassembly of the engine component 62. Further, the adjacent surfaces ofthe slots 86A-86D, 94A-94D and ribs 88A-88D, 96A-96D (e.g., the surfacesthat abut one another) provide joining interfaces for attachment (e.g.,by welding). That is, the slots 86A-86D, 94A-94D and ribs 88A-88D,96A-96D not only increase the ease of aligning the multiple pieces ofthe engine component 62, but also provide a joining interface forwelding.

While the illustrated example includes slots 86A-86D, 94A-94D and ribs88A-88D, 96A-96D, slots and ribs are not required in all embodiments.For instance, the radial supports 80A-80D may provide relatively planarmate faces corresponding to adjacent mate faces formed on the innersurfaces of the pressure side wall 70 and the suction side wall 72.These adjacent mate faces may provide joining interfaces for attachingthe pressure and suction side walls 70, 72 to the radial supports80A-80D.

As illustrated in FIG. 4, the pressure side wall 70 can include aplurality of microchannels 98, 100 configured to direct a cooling flowof fluid from a core passageway 102 provided between the pressure andsuction side walls 70, 72. The suction side wall 72 may likewise includea plurality of microchannels 104, 106 configured to direct a portion offluid from the core passageway 102 outward toward the suction side wall.The illustrated microchannels 98, 100, 104, 106 are shown forillustrative purposes only. The pressure and suction side walls 70, 72may include additional and/or different types of microchannels.

FIG. 5 is a view taken along line 5-5 from FIG. 2, and illustrates thedetail of the interaction between the slots in the pressure and suctionside walls 70, 72, and further illustrates the pressure and suction sidewalls 70, 72 joined to the radial supports 80A-80D. In one example, thepressure and suction side walls 70, 72 are welded (e.g., tack welded) tothe radial supports 80A-80D. In other examples, the pressure and suctionside walls 70, 72 are brazed, diffusion bonded, or glued/cemented to theradial supports 80A-80D. Other techniques for joining the pressure sidewall and suction side wall 70, 72 to the radial supports 80A-80D may beused herein.

FIG. 6 illustrates an example method for forming the engine component62. In the method, at 108, the pressure and suction side walls 70, 72are additively manufactured. As is known of additive manufacturingprocesses, a powdered metal may be selectively melted to form a firstlayer of the component. In one example, the powdered metal is a titaniumalloy. In another example, the powdered metal is a molybdenum alloy.Other example alloys may be used herein, such as: (1) a titanium, (2)tungsten, (3) nickel, (4) cobalt, and (5) steel alloys. As theselectively melted powder metal cools, additional powdered metal islayered on top of the cooled metal, and is selectively melted to formanother, subsequent layer. The process is repeated to build thecomponent. Example techniques include direct metal laser sintering(DMLS), electron beam melting (EBM), electron beam wire deposition(EBWD), and laser powder deposition (LPD). Other types of additivemanufacturing processes come within the scope of this disclosure. Forinstance, as illustrated in FIG. 7, one example additive manufacturingprocess includes the use of a laser powder plasma spray (LPPS) process.In this process, the external flowpath surfaces 70, 72 are additivelybuilt, and leverage the cast (or forged) radial supports 80A-80D as thebase for the additive build. In other words, the external flowpathsurfaces 70, 72 are essentially built onto the radial supports 80A-80D.

Separately, the root 66, the platform 68, and the radial supports80A-80D are formed, at 110, in one example by way of casting, such asinvestment casting. In another example, a forging process is used. Ineither case, the root 66, the platform 68, and the radial supports80A-80D may be formed of a single crystal, directionally solidified, orequiax alloy. Such alloys are generally more resistant to creep thanmaterials suited for additive manufacturing.

Finally, at 112, the pressure and suction side walls 70, 72 are joinedto the casting formed at 110, by way of welding, or some other joiningprocess, as mentioned above.

The microchannels 98, 100, 104, 106 in the pressure and suction sidewalls 70, 72 may be difficult to form by way of casting or forging. Onthe other hand, materials that are capable of being additivelymanufactured may be less resistant to creep and other stresses.Accordingly, this disclosure provides an engine component 62 with afirst portion (e.g., the root, platform, and radial supports 80A-80D)having sufficient creep and other load-resistive capabilities, andsecond and third portions (the pressure and suction side walls 70, 72,for example) including relatively intricate microchannel coolingpassageways for cooling the airfoil section 64.

While a molybdenum alloy is listed above as an example material foradditive manufacturing, the pressure and suction side walls 70, 72 maybe formed using different manufacturing techniques in the example whenthe pressure and suction side walls 70, 72 are made of a molybdenumalloy. For example, the pressure and suction side walls 70, 72 could beformed using EDM, ECM, or other, more conventional machining techniques.Further, if the pressure and suction side walls 70, 72 are made from amaterial that reacts with nickel (Ni) or has a different thermalexpansion coefficient, then an intermediate material coating may beapplied to the internal casting at locations of expected contact betweenthe pressure and suction side walls and the internal casting.

While the above discussion is made relative to a rotor blade, thisdisclosure extends to stator vanes as well. As illustrated in FIGS.8A-8B, an example stator vane 114 may include an inner platform 116, anouter platform 118, and a radial support 120 extending therebetween. Inthis example, the inner and outer platforms 116, 118 and the radialsupport 120 are formed using a casting or forging process. An airfoilsection 122 of the vane 114 includes pressure and suction side sections124A, 124B formed using an additive manufacturing process. Similar tothe above example, the pressure and suction side sections 124A, 124B mayinclude slots 126 for receiving a corresponding rib 128 from the radialsupport 120. In this example, the vane 114 only includes a single radialsupport 120, although it should be understood that additional radialsupports could be included.

Although the different examples have the specific components shown inthe illustrations, embodiments of this disclosure are not limited tothose particular combinations. It is possible to use some of thecomponents or features from one of the examples in combination withfeatures or components from another one of the examples.

One of ordinary skill in this art would understand that theabove-described embodiments are exemplary and non-limiting. That is,modifications of this disclosure would come within the scope of theclaims. Accordingly, the following claims should be studied to determinetheir true scope and content.

What is claimed is:
 1. A method of forming a component of a gas turbineengine, comprising: forming a first portion of the component using oneof a casting and a forging process, the first portion including a root,platform, and at least one radial support projecting from the platform;additively manufacturing a second portion of the component, the secondportion providing a side wall of the component; and joining the secondportion to the first portion; wherein the step of additivelymanufacturing the second portion includes directly additively buildingthe second portion onto the first portion, wherein the at least oneradial support of the first portion is used as a base for the additivebuild occurring during the step of additively manufacturing the secondportion such that the step of joining the second portion to the firstportion occurs concurrent with the step of additively manufacturing thesecond portion.
 2. The method as recited in claim 1, including:additively manufacturing a third portion of the component; and joiningthe third portion to the first portion.
 3. The method as recited inclaim 1, wherein the additive manufacturing step includes one of adirect metal laser sintering (DMLS) process, an electron beam melting(EBM) process, electron beam wire deposition (EBWD) process, a laserpowder deposition (LPD) process, and a laser powder plasma spray (LPPS)process.
 4. The method as recited in claim 1, wherein the additivemanufacturing step includes selectively melting a powdered metal, thepowdered metal being one of (1) a titanium alloy, (2) tungsten alloy,(3) nickel alloy, (4) cobalt alloy, (5) steel alloy, and (6) amolybdenum alloy.
 5. The method as recited in claim 2, wherein thesecond and third portions each include channels therein, the channelsformed during the steps of additively manufacturing the second and thirdportions, respectively, and configured to direct a cooling flow of fluidfrom a core passageway of the component toward an exterior of thecomponent.
 6. The method as recited in claim 5, wherein the secondportion is additively manufactured such that it includes a differentquantity of channels than the third portion.
 7. The method as recited inclaim 5, wherein the second portion is additively manufactured such thatit includes at least one different type of channel than the thirdportion.
 8. The method as recited in claim 2, wherein the second portionand the third portion provide pressure and suction side walls of thecomponent.
 9. The method as recited in claim 8, wherein the at least oneradial support includes at least one rib projecting into a correspondingslot formed in one of the second portion and the third portion.
 10. Themethod as recited in claim 1, wherein the component is one of a rotorblade and a stator vane.
 11. The method as recited in claim 1, whereinthe root, platform, and at least one radial support are integrallyformed of one of a single crystal, directionally solidified, and anequiax alloy.
 12. A method of forming a component of a gas turbineengine, comprising forming a first portion of the component using one ofa casting and a forging process, the first portion including a root,platform, and at least one radial support projecting from the platform;additively manufacturing a second portion of the component, the secondportion providing a side wall of the component; joining the secondportion to the first portion; additively manufacturing a third portionof the component; and joining the third portion to the first portion;wherein the second portion and the third portion provide pressure andsuction side walls of the component, wherein the at least one radialsupport includes at least one rib projecting into a corresponding slotformed in one of the second portion and the third portion, and whereinthe at least one radial support includes a plurality of radial supports,each of the radial supports including a first rib and a second ribprojecting into slots formed in the second portion and the thirdportion.
 13. The method as recited in claim 12, wherein the at least oneradial support provides a mate face corresponding to a mate face of oneof the second portion and the third portion.